The present invention relates generally to gas turbine engines, and, more specifically, to cooling of turbine rotor blades and stator vanes therein.
A gas turbine engine includes a compressor that compresses air which is channeled to a combustor wherein it is mixed with fuel and ignited for generating combustion gases. The combustion gases flow downstream through one or more stages of turbines which extract energy therefrom for powering the compressor and producing additional output power for driving a fan for powering an aircraft in flight for example.
A turbine stage includes a row of turbine rotor blades secured to the outer perimeter of a rotor disk, with a stationary turbine nozzle having a plurality of stator vanes disposed upstream therefrom. The combustion gases flow between the stator vanes and between the turbine blades for extracting energy to rotate the rotor disk.
Since the combustion gases are hot, the turbine vanes and blades are typically cooled with a portion of compressor air bled from the compressor for this purpose. Diverting any portion of the compressor air from use in the combustor necessarily decreases the overall efficiency of the engine. Accordingly, it is desired to cool the vanes and blades with as little compressor bleed air as possible.
Turbine vanes and blades include an airfoil over which the combustion gases flow. The airfoil typically includes one or more serpentine cooling passages therein through which the compressor bleed air is channeled for cooling the airfoil. The airfoil may include various turbulators therein for enhancing cooling effectiveness, and the cooling air is discharged from the passages through various film cooling holes disposed around the outer surface of the airfoil.
The airfoil outer surface is defined by a generally concave pressure side and an opposite, generally convex suction side which extend radially between a root and a tip of the airfoil and axially between leading and trailing edges thereof. The temperature profiles of the combustion gases channeled over the airfoil vary significantly over the pressure and suction sides. This in turn affects both the cooling requirements over the airfoil and cooling effectiveness. Greater cooling is desired where heat input is greatest, and backflow margin and blowing ratio must be controlled across the film cooling holes. Film cooling holes should have suitable blowing ratios to most effectively produce a protecting layer of film cooling air over the blade surface without flow separation and with suitable backflow margin.
A typical serpentine circuit includes internal ribs extending the full radial span of the airfoil from root to tip which alternately terminate adjacent thereto to define sequential channels or passes connected by 180.degree. reverse bends or turns. The cooling air increases in temperature in each pass as it removes heat from the airfoil and thusly creates an axial thermal gradient between the leading and trailing edges of the airfoil. In turn, thermal stress is created which affects the useful life of the airfoil.
In a turbine rotor blade, centrifugal loads must be carried through the airfoil to the rotor disk, and in turn effect substantial centrifugal stress in the blade. The radial span ribs of the serpentine circuits bridge the pressure and suction sides of the airfoil and are effective for carrying centrifugal loads. However, the interruption in the span ribs at the reverse turns do not carry centrifugal loads, and are typically located at the airfoil tip where such loads are smallest, and at the airfoil root where the airfoil is strongest.
These design constraints therefore limit the ability to optimize both airfoil cooling uniformity and centrifugal load carrying strength using typical serpentine circuits.
Accordingly, it is desired to provide an improved turbine airfoil for use in blades and vanes having more uniform cooling while maintaining airfoil strength.